WO2017085406A1

WO2017085406A1 - Propulsion unit comprising a main engine and an auxiliary engine

Info

Publication number: WO2017085406A1
Application number: PCT/FR2016/052979
Authority: WO
Inventors: Pascal Charles Edouard COAT; Jean-François Endy BERSOT; Stéphane ORCEL; Nicolas Jérôme Jean TANTOT
Original assignee: Safran Aircraft Engines
Priority date: 2015-11-16
Filing date: 2016-11-16
Publication date: 2017-05-26
Also published as: US20180327109A1

Abstract

The invention relates to a propulsion unit (2) configured to provide takeoff thrust and top of climb thrust and comprising: - a main engine (3), configured to supply main thrust during the takeoff phase and during the top of climb phase with a ratio (QT CHP ) between a temperature on the outlet side of the high-pressure compressor (34) in top of climb operating condition (T CHP(ToC)) and a temperature on the outlet side of the high-pressure compressor (34) in a takeoff operating condition (T CHP(TkOff)), which is comprised between 0.90 and 1.10, for example between 0.95 and 1.05, and - an auxiliary engine (4), distinct from the main engine (3) and configured to supply auxiliary thrust in order to supplement the main thrust of the main engine (3) at least during the takeoff phase.

Description

Propulsion unit comprising a main motor and an auxiliary motor

FIELD OF THE INVENTION

The invention relates to the general field of aircraft, and more particularly to the dimensioning of the engines of such aircraft with a view to improving, among other things, the specific consumption. The invention is applicable in all types of aircraft intended to perform missions with various operating conditions.

BACKGROUND

In operation, a given engine is requested differently according to the flight phases of the aircraft. Indeed, each phase of flight is associated with a condition of engine operation, including idle (or "idle" in English), takeoff (or "take off" in English), climb (or "climb" in English), the summit of climb (or "top of climb" or "maximum climb" in English) or the cruise (or "cruise" in English). During the aforementioned operating conditions, the engine is maintained for a relatively long time (between thirty seconds for takeoff and several hours for cruising) at predefined speed spectra, which depend on the engine redline (ie the absolute maximum speed encountered by the low pressure shaft during the entire flight). The most restrictive engine operating condition in terms of thrust is takeoff. This is why, usually, the engines for aircraft are dimensioned according to this operating condition in order to guarantee their ability to take off the aircraft. For this, the motors are dimensioned so as to operate at the maximum temperatures at the inlet and outlet of the combustion chamber during the take-off phase, so that the efficiency of the thermodynamic (and thus energy) cycle of the engine is optimal during this phase. These inlet and outlet temperatures of the combustion chamber will therefore directly affect the size of the high pressure parts of the engine (high pressure compressor, combustion chamber and high pressure turbine) and their constituent material, so that they are capable of to provide the thrust required for take-off.

However, the duration of the take-off phase is very short (between one and five minutes, depending on the type of aircraft and their mission) in front of the other phases of flight. As a result, during most of the flight, the engine requires a lower thrust and therefore has a thermodynamic efficiency (and therefore energy) less. This is particularly the case of the cruise operating condition, which generally lasts at least thirty minutes. Indeed, during the cruise, the power required by the engine is lower than during takeoff. However, the reduction of the power of the engine is obtained by reducing the temperature at the outlet of the combustion chamber and therefore at the inlet of the high-pressure turbine of the engine, which implies a reduction in the overall compression ratio. As a result, during this phase of flight, the specific consumption of the engine is greater than its optimum.

However, in order to comply with increasing regulatory constraints (in terms of acoustics and pollutant emissions in particular) and to reduce the operating costs of the engines, in particular related to their specific consumption, engine manufacturers tend to increase the temperature in particular. inlet and outlet of the combustion chambers to reduce the size of the high pressure body of the engines and increase the size of the low pressure body while maintaining acceptable fan diameters for aircraft manufacturers. Such an increase in the temperature at the inlet and at the outlet of the combustion chamber also makes it possible to improve the efficiency of the thermodynamic cycle of the engines, insofar as the overall compression ratio and the inlet temperature of the high-pressure turbine increase. This actually improves the thermodynamic efficiency during the takeoff phase, which is the dimensioning phase. However, the thermodynamic efficiency in the other phases of flight is not optimal, especially in cruising operating condition.

Engineers therefore seek to find a compromise between the needs of the engine following the different operating conditions and the impact of these constraints in terms of specific consumption, mass, acoustic constraints, etc. SUMMARY OF THE INVENTION

An object of the invention is therefore to propose a solution in the field of aircraft propulsion that responds to this problem of reconciliation of operational constraints, such as the ability of the propulsion unit to take off an aircraft, with objectives of ambitious fuel consumption, typical of civil commercial aviation.

For this purpose, the invention proposes a propulsion unit for an aircraft, said propulsion unit being configured to provide a takeoff thrust during a takeoff operating condition and a climb summit thrust during an operating condition. climbing summit and comprising:

at least one main engine, configured to provide a main thrust during the take-off operating condition and the climb-top operating condition, said main engine comprising a high-pressure compressor, and

- At least one auxiliary engine, separate from the main engine and configured to provide auxiliary thrust to complete the main thrust of the main engine during at least the takeoff operating condition.

Furthermore, the main motor is sized taking into account the thrust of the auxiliary motor in the operating condition of takeoff, such that a temperature ratio of the high pressure compressor, corresponding to the ratio between a temperature at the output of the high pressure compressor of the main engine in the operating condition of the ascent summit and a temperature at the outlet of the high pressure compressor of the engine main operating condition of takeoff, between 0.90 and 1 .10, for example between 0.95 and 1 .05.

Independently or in combination, the main motor may comprise a streamlined blower having an inlet section, said blower being located upstream of the high pressure compressor in the direction of gas flow in the main engine. A reduced main engine blower speed ratio, corresponding to the ratio of the reduced air flow entering the main engine blower at the input section to the climb top operating condition and the reduced flow rate of the engine. The air entering the main engine blower at said entry section in take-off operating condition may be between 1.3 and 1.5, preferably between 1.35 and 1.40. Also independently or in combination, the main motor may have a ratio of between 1.50 and 1.90, for example between 1.55 and 1.80, between its overall compression ratio in the climb-up condition and its overall compression ratio of the main engine in take-off operating condition.

Also independently or in combination with the above features, the main engine may further include a combustion chamber extending downstream of the high pressure compressor in the direction of gas flow in the main engine. The main engine can then have a temperature ratio (corresponding to the ratio between, on the one hand, a ratio between a temperature at the outlet of the combustion chamber of the main engine in peak operating condition of a rise and a temperature at the outlet of the combustion chamber of the main engine in take-off operating condition, and secondly, the temperature ratio of the high-pressure compressor) of between 1 .00 and 1 .10.

Also independently or in combination, the main motor may have a body size ratio (corresponding to the ratio of a body size at an inlet section of the main engine high pressure compressor to a climb top condition of operation and the body size at said inlet section of the main engine high pressure compressor in take-off operating condition of between 0.95 and 1.05.

Still independently or in combination, the main engine may comprise, downstream of the blower, a combustion chamber in the direction of gas flow, and a ratio between a temperature at the outlet of the combustion chamber of the main engine under the condition of climb top operation and a temperature at the outlet of the combustion chamber of the main engine in take-off operating condition can then be between 0.90 and 1 .10, for example between 1 .00 and 1 .05.

Also independently or in combination, the main engine may comprise a high pressure turbine downstream of the high pressure compressor in the gas flow direction, with a ratio between a temperature output of the high pressure turbine of the main engine in operating condition and a temperature at the outlet of the main engine high pressure turbine in take-off operating condition which is between 0.90 and 1 .10, for example between 0.95 and 1 .05. Also independently or in combination, the auxiliary motor comprises a ducted blower having an inlet section, with a reduced blower speed ratio of the auxiliary engine, corresponding to the ratio of the reduced air flow entering the blower of the auxiliary engine to the level of the inlet section in the up-hill operating condition and the reduced flow rate of air entering the auxiliary engine blower at said take-off operating condition input section, between 1 .00 and 1 .10. Also independently or in combination, a ratio between an overall compression ratio of the auxiliary engine in climb-up operating condition and an overall compression ratio of the auxiliary engine in take-off operating condition can be between 1 .00 and 1. 30.

Finally, the propulsion unit may, in a nonlimiting manner, comprise at least two auxiliary engines, the thrust of said auxiliary engines participating to 100% of the auxiliary thrust. BRIEF DESCRIPTION OF THE DRAWINGS

Other features, objects and advantages of the present invention will appear better on reading the detailed description which follows, and with reference to the appended drawings given by way of non-limiting examples and in which:

FIG. 1 is a graph illustrating, for several parameters, the ratio between the value of this measured parameter for an operating condition corresponding to the climb summit and the value of this parameter measured for an operating condition corresponding to the take-off, for an example embodiment of a main engine of a propulsion unit according to the invention and for a conventional engine, FIG. 2 illustrates an exemplary embodiment of an aircraft that may comprise a propulsion unit according to the invention, and Figure 3 is a schematic partial sectional view of an exemplary embodiment of a main motor.

DETAILED DESCRIPTION OF AN EMBODIMENT

In order to improve the specific consumption of a propulsion unit

2 for an aircraft 1 comprising a main engine 3, the invention proposes to release the main engine 3 from the constraint of being able to provide sufficient thrust to take off the aircraft 1 and to add to the propulsion unit 2 an auxiliary motor 4, separate from the main engine 3, in order to compensate for the thrust loss associated with this modification of the main engine 3. It then becomes possible to size the main engine 3 by significantly improving its specific fuel consumption in the flight phases having a important duration, as the cruise, while ensuring that the propulsion unit 2 is capable of taking off the aircraft 1.

For this, the propulsion unit 2 is configured to operate at at least two different operating conditions and comprises at least one main engine 3 and an auxiliary engine 4. These two engines contribute to the total thrust delivered by the propulsion unit, in different thrust proportions depending on the flight phase. By main engine means here and throughout the present text a motor configured to provide thrust during all the different phases of flight and in particular to provide during the cruise phase a thrust that contributes primarily to the total thrust . Auxiliary engine means a motor that assists the main engine by providing auxiliary thrust during certain phases of flight (during the take-off phase and up to the climb summit, in particular). Preferably, the auxiliary engine is cut during flight phases requiring less total thrust, such as the cruise phase; during these phases, it can also operate at low speed or at low thrust. In what follows, the invention will more particularly be described in the case where the main engine 3 comprises a turbojet engine. However, this is not limiting, the main engine (s) 3 possibly comprising one or more turbojet engines and / or one or more turboprop engines, said main engines 3 possibly comprising at least one ducted or non-ducted fan / propeller.

In a manner known per se, the turbojet engine 3 thus comprises, from upstream to downstream in the direction of flow of the gases in the turbojet engine 3, at least one ducted fan 30 housed in a fan casing 30, an annular space of primary flow and an annular secondary flow space. The mass of air sucked by the fan 30 is thus divided into a primary flow, which circulates in the primary flow space, and a secondary flow, which is concentric with the primary flow and circulates in the space of secondary flow.

The primary flow space passes through a primary body comprising one or more stages of compressors, for example a low pressure compressor 32 and a high pressure compressor 34, a combustion chamber 36, one or more turbine stages, for example a turbine high pressure 38 and a low pressure turbine 40, and a gas exhaust nozzle.

According to the flight phases, the main engine 3 and the auxiliary motor 4 together provide the thrust of the propulsion unit. In particular, the main engine 3 may be assisted by the auxiliary engine 4 in the take-off phase to provide the take-off thrust to the propulsion unit 2 and possibly in the climb summit phase to provide the climb summit thrust. For example, the thrust provided by the propulsion unit 2 during the take-off phase can be obtained up to 5% to 45% by the auxiliary engine 4, the complement being provided by the main engine 3. In the climb summit phase , the main motor 3 can provide all the necessary thrust, or be assisted from 0% to 50% by the auxiliary motor 4. Typically, for an engine having a low speed rotational speed reduction of between 3,000 rpm (revolutions per minute) and 4000 rpm, the takeoff corresponds to a rotation speed of the low pressure shaft of between 2500 and 3000 rpm, while the climbing top corresponds to a rotational speed of the low pressure shaft of between 3000 rpm and 3500 rpm. Furthermore, the propulsion unit 2 may have additional operating conditions, such as, among other things, cruising, idling (on the ground and in flight), etc.

It will be noted that the distribution of the thrust between the main engine 3 and the auxiliary engine 4 of the propulsion unit 2 can be determined according to the type of aircraft 1 and the associated mission type (short, medium, long haul, etc. .). Typically, for an aircraft 1 configured to perform a mission of the long-haul type, the share of the thrust provided by the auxiliary engine 4 at the climbing summit is preferably greater than in the case of an aircraft 1 configured to perform a short-haul type mission. Indeed, the flight time in cruising operating condition is shorter on a short-haul than on a long haul, so it may be preferable to improve the thermodynamic efficiency of the propulsion unit 2 at the top. to increase and limit the size and weight of the auxiliary engine 4 rather than to improve its thermodynamic efficiency in cruising and to increase the size and weight of the auxiliary engine 4.

The auxiliary engine 4 may provide a continuous thrust between the operating condition corresponding to the take-off and the operating condition corresponding to the climbing summit, or alternatively be stopped during at least one of said speeds.

In order to reduce the specific consumption of the propulsion unit 2 while guaranteeing the capacity of the propulsion unit 2 to take off an aircraft 1, the main engine 3 is dimensioned so that a temperature ratio of the high pressure compressor QTCHP is between 0.90 and 1.10, for example between 0.95 and 1.05. This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

In relation to the temperature of the high-pressure compressor QT _C HP _, the ratio between the temperature at the outlet of the high-pressure compressor 34 (and therefore at the combustion chamber inlet 36) of the main engine 3 under the operating condition of the ascent vertex will be understood here. TCHP (T _O C) and a temperature at the outlet of the high pressure compressor 34 of the main engine 3 under take-off operating condition T _C HP (Tkoff) - The output temperature of the high pressure compressor T _C HP represents the temperature of the fluid at the outlet of the diffuser, which is itself placed behind the last mobile wheel of the high pressure compressor 34.

By way of comparison, for a conventional engine (that is to say a motor sized from the take-off operating condition and which has no auxiliary motor), the QT _C HP temperature ratio is generally between 0.85. and 0.95. It can be deduced from this that the temperature T _C HP at the outlet of the high pressure compressor 34 at the summit of rise is higher in the main engine 3 than in a conventional engine. The main engine 3 is dimensioned so as to have very few variations of T _C HP at the compressor outlet, between the take-off condition and the climb crown condition, compared with a conventional engine without auxiliary thrust at take-off. The compression ratio of the high pressure compressor 34 is therefore higher for the main engine 3 at the top of the climb, which is a benefit in terms of the thermal efficiency of the turbojet (s) / turboprop (s) of the main engine 3.

In a turbojet type engine, the air pressure at the outlet of the high pressure compressor 34 is the highest of the engine. As a result, the high pressure compressor 34 can not be cooled, since none of the other components is likely to provide it with enough pressurized air to ventilate it. The temperature at the outlet of the high pressure compressor 34 is therefore an optimization point of this compressor. By dimensioning the main engine 3 so that the temperature T _C HP (TOC) at the summit of rise is greater than the TcHP temperature (Tkoff) at takeoff, it is thus possible to optimize the thermodynamic cycle of the turbojet of the main engine 3 in climb-up or cruising operating condition, instead of having a compromise between optimizations in climb-up operating condition and take-off operating condition, and improve the main-engine specific power consumption. that, knowing the optimum temperature T HP _C to reach the output of the high pressure compressor 34, it is then possible to define an optimum shape of the blades of each stage of the high pressure compressor 34, associated with a material technology.

In the case where the main engine 3 comprises at least one turbojet engine, a reduced fan _fan speed ratio of the main engine 3 may be between 1.30 and 1.50. Compared with the reduced fan flow rate Q _fan of the main engine 3, here the ratio between the reduced flow of air entering the fan of the main engine 3 at the inlet section in the operating condition of the top of the engine will be understood. mounted and the reduced flow of air entering the blower 30 of the main engine 3 at said inlet section in take-off operating condition. The reduced flow rate Q _fan corresponds here to the total mass air flow at the inlet of the fan Qm _fan and reduced with the total pressure and temperature conditions at the inlet of the fan according to the following formula:

Qfan - Q ^m fan

where: - Qm _f an corresponds to the total mass air flow at the intake section of the blower - Tfan is the temperature at the inlet section of the blower (expressed in Kelvin, K)

- T _s td corresponds to the standard temperature (288.15 K)

- Pfan corresponds to the pressure at the inlet section of the fan (expressed in Bar)

- P _s td corresponds to the standard pressure (1 .0135 Bar)

The inlet section of the blower 30, where the Qnrifan air flow rate, the T _fan temperature and the _fan P pressure are measured, corresponds to the area of the blower housing 30 as seen by the flow entering said blower 30. , in a plane perpendicular to an axis of revolution of the fan 30. It will be noted that the exact position of the measurement of this input section is not decisive insofar as a flow ratio is evaluated, as long as the flow rate is determined for the same inlet section of the blower 30 in take-off operating condition and in climb-up operating condition.

To calculate this _fan ratio Q, the reduced air flow rate in a climb-up condition and in take-off operating condition is measured when the main engine 3 is stationary in a standard atmosphere (as defined by the engine manual). International Civil Aviation Organization (ICAO), Doc 7488/3, 3rd Edition) and at sea level.

A main engine 3 comprising a turbojet having such a reduced fan speed ratio Qf _an then has a better specific consumption compared to a conventional engine since it is dimensioned not according to a compromise between the operating condition of takeoff and the cruising operating condition, but mainly depending on the climb and cruise summit operating condition, which correspond to a substantial part of the operation of the main engine 3. The reduced airflow Q _fan at the input of the blower 30 of this main motor 3 is therefore more important at the summit of climb than at takeoff whereas, for a conventional engine, the ratio of reduced flow rate of fan Qf _year is between 1 .00 and 1 .10. As a result, a main motor 3 according to the invention has a more efficient thermodynamic cycle than a conventional motor.

In one embodiment, the total pressure ratio of the blower 30 of the main engine 3 may be between 1.35 and 1.40. The ratio Q _OPR between the overall compression ratio of the main engine 3 in climb-up operating condition and the overall compression ratio of the main engine 3 under take-off operating condition can be between 1.50 and 1.90, for example between 1.55 and 1.80. This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

For comparison, for a conventional engine, this ratio is usually between 1 .00 and 1 .30. This difference is explained by the fact that the assistance of the auxiliary engine 4 optimizes the thermodynamic operation of the main engine 3 by choosing by design to operate it for all operating conditions (takeoff, climb summit, cruise, idle, etc.) at temperatures and pressures close to the maximum permitted by the nature of the materials and components of its modules. This makes it possible in particular to increase the compression ratio in the low pressure and high pressure compressors of the main engine 3.

By overall compression ratio, here will be understood the combination of the compression ratio of the high pressure compressor 34, the low pressure compressor 32 and the blower 30 or, in other words, the ratio between the outlet pressure of the high pressure compressor. 34 (and thus at the combustion chamber inlet 36) and the pressure at the inlet of the fan 30. The overall compression ratio is determined, whether in run-up operating condition or in take-off operating condition, when the main engine 3 is stationary in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) manual, Doc 7488 / 3, 3rd edition) and at sea level.

The temperature ratio QT _Co mb, corresponding to the ratio between the temperature at the outlet of the combustion chamber 36 (and thus at the inlet of the high-pressure turbine 38) of the main engine 3 under the operating condition of the climb vertex T _Co mb ( Toc) and the temperature at the outlet of the combustion chamber 36 of the main engine 3 in the take-off operating condition T _Co mb (Tkoff) can be between 0.90 and 1 .10, for example between 0.95 and 1 .05. This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

By way of comparison, for a conventional engine, the QT _Co mb temperature ratio is generally between 0.85 and 0.95. It can be deduced that the temperature T _Co mb at the outlet of the combustion chamber 36 at the summit of rise is higher in the main engine 3 than in a conventional engine. The thermodynamic cycle of the turbojet of the main engine 3 is therefore more efficient.

In a turbojet type engine, the high-pressure turbine 38 is generally cooled by ventilation. The design of the cooling system is generally performed on the maximum temperature conditions encountered at the take-off condition, and the cooling system is oversized. and underused for other operating conditions. The QTcomb temperature ratio thus defined makes it possible to continuously use the cooling system of the high-pressure turbine 38 of the main engine 3 over its optimum operation and therefore cooling efficiency. In addition, the limitation of the thermal excursions seen by the high-pressure turbine 38 between take-off and cruise conditions contributes to limiting the mechanical degradation of the latter and thus to improve its service life.

A high pressure temperature ratio QT _Co mb / QTcHP, corresponding to the ratio between, on the one hand, the ratio between the temperature at the outlet of the combustion chamber 36 of the main engine 3 in the operating condition of the climb vertex T _Co mb (Toc) and the temperature at the outlet of the combustion chamber 36 of the main engine 3 in the take-off operating condition T _Co mb (Tkoff), and secondly, the ratio QT _C HP between the outlet temperature of the high compressor pressure 34 of the main engine 3 in the up-hill operating condition T _C HP (T _O C) and a temperature at the outlet of the high-pressure compressor 34 of the main engine 3 under the take-off operating condition T _C HP (Tkoff), can be between 1 .00 and 1 .10.

In other words, the high pressure temperature ratio

QTcomb QTcHP is the ratio of the QT _Co mb temperature ratio to the QT _C HP-temperature ratio.

This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

It will therefore be understood that here, the main engine 3 is not a variable cycle engine, since its high pressure temperature ratio QTcomb QTcHP is substantially equal to that of a conventional engine regardless of the operating conditions. The temperature ratio QT _T HP, which corresponds to the ratio between the outlet temperature of the high-pressure turbine 38 (and thus at the inlet of the low-pressure turbine 40) of the main engine 3 in the operating condition of the climb vertex T _T HP (T _OC ) and the outlet temperature of the high-pressure turbine 38 of the main engine 3 in the take-off operating condition T _T HP (Tkoff) can be between 0.90 and 1.10, for example between 0.95 and 1.05. The temperature at the outlet of the high pressure turbine T _T HP may, for example, be measured in an area close to the last impeller of the high-pressure turbine 38 (at the leading edge of the first distributor of the low-pressure turbine 40 or at the intrados wall of the second distributor of the low-pressure turbine 40). This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

By way of comparison, for a conventional engine, the QTTHP temperature ratio is generally between 0.85 and 0.95. It can be deduced that the temperature at the outlet of the low-pressure turbine 40 at the summit of rise is higher in the main engine 3 than in a conventional engine.

The inlet temperature of the low pressure turbine 40 is an optimization point of the low pressure turbine 40 and the main motor 3 in general. The choice of the outlet temperature of the high-pressure turbine 38 under the T _T HP (T _OC ) up-hill operating condition thus makes it possible to dimension the main engine 3 in the operating condition of the ascent or crest, which cover a substantial part of the operation of the main engine 3, and not exclusively in take-off operating condition. The limitation of the thermal excursions seen by the low-pressure turbine 40 between the take-off and cruising conditions contributes to limiting the mechanical degradation of the latter and thus to improving its service life.

A C-body size ratio of the main engine 3 between the climb-up and take-off operating conditions can be between 0.95 and 1.05. This relationship is valid regardless of the type of the main engine 3 (or a number of turbojet (s) and / or turboprop (s)).

Relative body Q _∞ re size of the main engine 3, it is understood here the ratio between the body size at a high pressure compressor inlet section 34 of the main motor 3 in mounted condition of the top of operation and the body size at level of said input section in take-off operating condition.

The size of body T _∞ re here corresponds to the air mass flow rate Qm _CO re entering the high pressure compressor 34 of the main engine 3 at the corrected input section TCHP temperature conditions and pressure P _C Total HP at the outlet of the high pressure compressor 34 according to the following formula:

where: - QITICHP corresponds to the total mass air flow at the inlet of the blower

- TCHP corresponds to the temperature at the outlet of the high-pressure compressor 34 (expressed in Kelvin, K)

- T _s td corresponds to the standard temperature (288.15 K)

- PCHP corresponds to the pressure at the outlet of the high pressure compressor 34 (expressed in Bar)

- P _s td corresponds to the standard pressure (1 .0135 Bar)

Here again, the body size T _CO re in climb-up operating condition and in take-off operating condition is measured when the main engine 3 is stationary in a standard atmosphere (as defined by the manual of the Organization). International Civil Aviation (ICAO), Doc 7488/3, 3rd Edition) and at sea level.

The body size T _∞ re is representative of the geometric height of the vein of the high pressure compressor 34.

The auxiliary motor 4 may also be dimensioned so as to optimize the specific consumption of the propulsion unit 2. Typically, when the auxiliary engine 4 comprises one or more turbojets having, in a conventional manner, a ducted fan, a ratio of blower rate Q _fan of the auxiliary motor 4 can be between 1 .00 and 1 .10. In a similar manner to the fan flow ratio Qf _an of the main motor 3 defined above, the fan flow rate ratio Qf _an of the auxiliary engine 4 then corresponds to the ratio of the air flow rate entering the fan 30 of the auxiliary engine 4 at the input section in upwind operating condition and the air flow entering the blower 30 of the auxiliary engine 4 at said input section in take-off operating condition, the flow being measured when auxiliary engine 4 is stationary in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) Manual, Doc 7488/3, 3rd Edition) and at sea level.

Alternatively, the auxiliary engine (s) 4 may comprise one or more turboprop engines and / or one or more propulsive effectors driven by electric motors. According to another variant, the auxiliary engine or engines may comprise one or more turbojet engines in combination with one or more turboprop engines and / or one or more propulsive effectors driven by electric motors.

Moreover, the ratio Q _OPR between the overall compression ratio of the auxiliary engine 4 in the climb-summit operating condition and the overall compression ratio of the auxiliary engine 4 in take-off operating condition can be between 1 .00 and 1. .30.

Here again, the overall compression ratio in climb-up operating condition and take-off operating condition is measured when the auxiliary engine 4 is stationary in a standard atmosphere (as defined by the Organization's manual of the international civil Aviation Organization (ICAO), Doc 7488/3, 3rd edition) and at the sea. A ratio of body size Q _∞ re auxiliary motor 4 between the conditions of climb and top off operation can be understood between 0.95 and 1.05. In relation to the body size Q _∞ of the auxiliary motor 4, here the ratio between the body size at an inlet section of the high pressure compressor 34 of the auxiliary motor 4 in the operating condition of the ascent vertex will be understood. and the body size at said input section in take-off operating condition.

The definition and measurement of body size T _∞ re specified for the main motor 3 applies mutatis mutandis to the auxiliary motor 4. The propulsion system 2 may comprise one or more main engines 3 and one or more auxiliary motors 4. In this case, the main engine (s) 3 then participate together in the supply of the main thrust, while the auxiliary engine (s) 4 participate together in the supply of the auxiliary thrust.

For example, the propulsion unit 2 may comprise a main engine 3 and two auxiliary engines 4. The auxiliary engines 4 may for example be fixed under the wings of an aircraft 1 while the main engine 3 may be placed at the rear of the fuselage of the aircraft 1, as illustrated in FIG. 2.

Typically, the propulsion unit 2 may comprise a turbofan propeller propeller and two auxiliary engines 4 each comprising one or more effectors driven by an electric motor.

If necessary, the auxiliary engine (s) 4 may be retractable, that is to say that their position may be modified during certain phases of the flight of the aircraft 1 in order to minimize their drag. For example, the auxiliary engines 4 can be retracted by being tucked into a specific wedge formed in the wings of the aircraft 1.

Claims

1. A propulsion assembly (2) for an aircraft (1), said propulsion assembly (2) being configured to provide a take-off thrust during a take-off operating condition and a climb-up thrust during a take-off condition. summit climb operation and comprising:

at least one main engine (3), configured to provide a main thrust during the take-off operating condition and the climb-top operating condition, and

at least one auxiliary motor (4), distinct from the main engine (3) and configured to provide auxiliary thrust to complete the main thrust of the main engine (3) during at least the take-off operating condition,

the propulsion unit (2) being characterized in that:

the main motor comprises a high-pressure compressor (34), and

the main motor is dimensioned taking into account the thrust of the auxiliary engine in the take-off operating condition, such that a temperature ratio of the high-pressure compressor (QTCHP) corresponding to the ratio between a temperature at the outlet of the compressor high pressure (34) of the main engine (3) in up-hill operating condition (T _C HP (TOC)) and a temperature at the outlet of the high-pressure compressor (34) of the main engine (3) in operating condition of takeoff (TcHP (Tkoff)), being between 0.90 and 1 .10, for example between 0.95 and 1 .05.

2. propulsion unit (2) according to claim 1, wherein:

the main motor (3) further comprises a ducted fan (30) having an inlet section, said fan (30) being situated upstream of the high pressure compressor (34) in the direction of flow of the gases in the main motor (3), and a reduced fan speed ratio (Qf _an ) of the main engine (3) corresponding to the ratio of the reduced air flow entering the blower (30) of the main engine (3) at the inlet section; in the up-hill operating condition and the reduced flow rate of air entering the main engine blower (30) at said entry section in take-off operating condition is between 1 .3 and 1.50, preferably between 1.35 and 1.40.

3. propulsion unit (2) according to one of claims 1 or 2, wherein a ratio (QOPR) between an overall compression ratio of the main engine (3) in upwind operating condition and a global compression ratio the main engine (3) in take-off operating condition is between 1.50 and 1.90, for example between 1.55 and 1.80.

4. propulsion unit (2) according to one of claims 1 to 3, wherein:

the main motor (3) further comprises a combustion chamber (36) extending downstream of the high pressure compressor (34) in the direction of gas flow in the main engine (3), and

a temperature ratio (QT _Co mb / QTcHp) of the main engine (3), corresponding to the ratio between, on the one hand, a ratio between a temperature at the outlet of the combustion chamber (36) of the main engine (3) in a climb summit operating condition (T _Co mb (Toc)) and a temperature at the outlet of the combustion chamber (36) of the main engine (3) under take-off operating condition (T _Co mb (Tkoff)), and secondly, the ratio (QT _C HP) of the high pressure compressor temperature, is between 1 .00 and 1 .1 0.

5. Propulsion unit (2) according to one of claims 1 to 4, wherein a body size ratio (Q∞re) of the main motor (3), corresponding to the ratio between a body size at a level of section input of high pressure compressor (34) of main engine (3) in upwind operating condition and body size at said inlet section of high pressure compressor (34) of main engine (3) in take-off operating condition, is between 0.95 and 1 .05.

6. propulsion unit (2) according to one of claims 1 to 5, wherein:

the main motor (3) further comprises, downstream of the fan (30), a combustion chamber (36) in the direction of flow of the gases in the main engine (3), and

a ratio (QT _Co mb) between a temperature at the outlet of the combustion chamber (36) of the main engine (3) in a climb summit operating condition (T _Co mb (Toc)) and a temperature at the outlet of the combustion chamber (36) of the main engine (3) in take-off operating condition (T _Co mb (Tkoff)) is between 0.90 and 1 .10, for example between 1 .00 and 1 .05.

7. propulsion unit (2) according to one of claims 1 to 6, wherein:

the main motor (3) further comprises, downstream of the high-pressure compressor (34) in the direction of flow of the gases in the main motor (3), a high-pressure turbine (38), and

a ratio (QT _T HP) between a temperature at the outlet of the high-pressure turbine (38) of the main engine (3) in a climb-summit operating condition (T _T HP (TOC)) and a temperature at the outlet of the high pressure turbine (38) of the main engine (3) in take-off operating condition (T _T HP (Tkoff)) is between 0.90 and 1 .10, for example between 0.95 and 1 .05.

8. propulsion unit (2) according to one of claims 1 to 7, wherein the auxiliary motor (4) comprises a fan (30) streamlined having an inlet section, and wherein a reduced auxiliary motor blower speed ratio (4), corresponding to the ratio of the reduced air flow entering the blower (30) of the auxiliary engine (4) to the level of the input section in the up-hill operating condition and the reduced flow rate of air entering the blower (30) of the auxiliary motor (4) at said input section under take-off operating condition is included between 1 .00 and 1 .10.

9. Propulsion unit (2) according to one of claims 1 to 8, wherein a ratio between an overall compression ratio of the auxiliary motor (4) in up-hill operating condition and an overall compression ratio of the auxiliary motor. (4) in take-off operating condition, is between 1 .00 and 1 .30.

10. Propulsion unit according to one of claims 1 to 9, comprising at least two auxiliary motors (4), the thrust of said auxiliary motors (4) participating to 100% of the auxiliary thrust.